Method and apparatus to enhance laminar flow for gas turbine engine components

ABSTRACT

A gas turbine engine component has a component body configured to be positioned within a flow path of a gas turbine engine having an external pressure, and wherein the component body includes at least one internal cavity having an internal pressure. At least one inlet opening is formed in an outer surface of the component body to direct hot exhaust gas flow into the at least one internal cavity, and there is at least one outlet from the internal cavity. The internal pressure is less than an inlet external pressure at the inlet opening and the internal pressure is greater than an outlet external pressure at the outlet opening to controllably ingest hot exhaust gas via the inlet opening and expel the hot exhaust gas via the outlet opening to maintain a laminar boundary layer along the outer surface of the component body.

RELATED APPLICATION

This application is a continuation-in-part of U.S. application Ser. No.15/155,146, filed May 16, 2016.

BACKGROUND OF THE INVENTION

Gas turbine engines are continually being driven to provide higherthrust efficiencies to maximize performance. As hot engine gases flowaround airfoils in the turbine section, turbulent flow can be generatedalong an external surface of the airfoils, which adversely affectsefficiencies. One proven method to improve efficiency includesoptimizing airfoil/endwall shapes and minimizing turbine cooling airusage to reduce friction and pressure drag around the airfoil.

Airfoil friction drag is created by a combination of friction loss andmixing loss. Friction loss from near-wall turbulence that impacts aboundary layer along the external surface of the airfoil can result in asignificant energy loss. In a cooled airfoil configuration, film coolingair may be targeted towards different areas or via different shapedholes to lessen the impact on friction and mixing loss factors. In anuncooled airfoil configuration, such features are not available makingit difficult to improve efficiency.

SUMMARY OF THE INVENTION

In a featured embodiment, a gas turbine engine component has a componentbody configured to be positioned within a flow path of a gas turbineengine having an external pressure, wherein the component body includesat least one internal cavity having an internal pressure. At least oneinlet opening is formed in an outer surface of the component body todirect hot exhaust gas flow into the at least one internal cavity. Atleast one outlet from the internal cavity, wherein the internal pressureis less than an inlet external pressure at the inlet opening and theinternal pressure is greater than an outlet external pressure at theoutlet opening to controllably ingest hot exhaust gas via the inletopening and expel the hot exhaust gas via the outlet opening to maintaina laminar boundary layer along the outer surface of the component body.

In another embodiment according to the previous embodiment, thecomponent body comprises at least one of an airfoil in a turbine orcompressor, a platform, or a transition duct.

In another embodiment according to any of the previous embodiments, thecomponent body comprises an airfoil having a leading edge, a trailingedge, and pressure and suction side walls extending from the leadingedge to the trailing edge, and wherein the airfoil extends from a baseto a tip.

In another embodiment according to any of the previous embodiments, atleast one inlet opening comprises a plurality of inlet openings formedin one or both of the pressure and suction side walls.

In another embodiment according to any of the previous embodiments, theleading edge is free from inlet openings.

In another embodiment according to any of the previous embodiments, atleast one outlet comprises at least one opening to the external surfacethat is located near or at the trailing edge.

In another embodiment according to any of the previous embodiments, atleast one outlet comprises at least one opening to the external surfacethat is located near or at the tip.

In another embodiment according to any of the previous embodiments, thecomponent body is a non-cooled component and the internal cavity is freefrom receiving cooling flow.

In another embodiment according to any of the previous embodiments, atleast one inlet opening provides a passage surface that is coated withat least one of a thermal barrier coating or environmental barriercoating.

In another embodiment according to any of the previous embodiments, theinternal cavity is coated with at least one of a thermal barrier coatingor environmental barrier coating.

In another embodiment according to any of the previous embodiments, theinternal cavity and the at least one inlet opening are coated with aplurality of coatings including at least one thermal barrier coating andat least one environmental barrier coating.

In another embodiment according to any of the previous embodiments, thecomponent body is comprised of a non-metallic material.

In another embodiment according to any of the previous embodiments, thenon-metallic material is a ceramic matrix composite (CMC) material.

In another featured embodiment, a gas turbine engine has a componentbody configured to be positioned within a flow path of a gas turbineengine having an external pressure, wherein the component body includesat least one internal cavity having an internal pressure. A plurality ofinlet openings formed in an external surface of the component body todirect hot exhaust gas flow into the at least one internal cavity,wherein the inlet openings include a passage surface that is coated withat least one of a thermal barrier coating or environmental barriercoating. At least one outlet from the internal cavity, wherein theinternal pressure is less than an inlet external pressure at the inletopening and the internal pressure is greater than an outlet externalpressure at the outlet opening to controllably ingest hot exhaust gasvia the inlet opening and expel the hot exhaust gas via the outletopening to maintain a laminar boundary layer along the outer surface ofthe component body.

In another embodiment according to the previous embodiment, thecomponent body comprises an airfoil having a leading edge, a trailingedge, and pressure and suction side walls extending from the leadingedge to the trailing edge, and wherein the airfoil includes the at leastone internal cavity which is free from receiving cooling air flow, andwherein the plurality of inlet openings are formed within at least oneof the pressure and second sides, and wherein the leading edge comprisesa showerhead region that is free from holes or slots, and wherein the atleast one outlet comprises at least one opening to the external surfacethat is located near or at the trailing edge.

In another embodiment according to any of the previous embodiments, acompressor section is included. A combustor section is downstream of thecompressor section. A turbine section is downstream of the combustorsection, and includes at least a high pressure turbine downstream of thecombustor section and a low pressure turbine downstream of the highpressure turbine, and wherein the airfoil is located within the lowpressure turbine.

In another embodiment according to any of the previous embodiments, theinternal cavity is coated with at least one of a thermal barrier coatingor environmental barrier coating.

In another embodiment according to any of the previous embodiments, theinternal cavity and the at least one inlet opening are coated with aplurality of coatings including at least one layer of a thermal barriercoating and at least one layer of an environmental barrier coating.

In another embodiment according to any of the previous embodiments, thecomponent is made of a non-metallic material comprising a CMC material.

In another featured embodiment, a method of enhancing laminar flow for agas turbine engine component includes the a step of positioning acomponent body within a hot gas flow of a gas turbine engine having anexternal pressure. The component body includes at least one internalcavity having an internal pressure. At least one inlet opening isprovided in an external surface of the component body to direct hotexhaust gas flow into the at least one internal cavity. The method alsoincludes the step of providing at least one outlet from the internalcavity to external atmosphere. The internal pressure is maintained to beless than the external pressure at the inlet opening and to be greaterthan the external pressure at the outlet opening to controllably ingesta portion of the hot gas flow via the inlet opening and expel ingestedhot exhaust gas via the outlet opening to maintain a laminar boundarylayer of a remaining portion of the hot gas flow along the outer surfaceof the component body.

In another embodiment according to the previous embodiments, thecomponent body comprises one of an airfoil, a platform or a transitionduct in at least one of a mid-turbine frame or turbine exhaust case andincluding keeping the internal cavity free from cooling flow.

In another embodiment according to any of the previous embodiments, thecomponent body is formed from a non-metallic material. At least one ofthe internal cavity and the at least one inlet opening are coated withat least one of a thermal barrier coating or environmental barriercoating.

The foregoing features and elements may be combined in any combinationwithout exclusivity, unless expressly indicated otherwise.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic representation of one example of a gas turbineengine.

FIG. 2 is a schematic representation of a top of an airfoilincorporating the subject invention.

FIG. 3 is a side view of the airfoil of FIG. 2.

FIG. 4 is graph depicting an amount of turbulent flow generated as flowmoves from a leading edge to trailing edge of an airfoil that does notinclude the subject invention.

FIG. 5 is graph depicting an amount of turbulent flow generated as flowmoves from a leading edge to trailing edge of an airfoil that doesinclude the subject invention.

FIG. 6 is a schematic representation of a transition duct incorporatingthe subject invention.

FIG. 7 is a schematic representation of a platform incorporating thesubject invention.

FIG. 8 is a schematic representation of a section of an airfoilincorporating another example of the subject invention.

FIG. 9 is a schematic representation of a gas turbine engine componentincluding a coating layer.

FIG. 10 is a schematic representation of using multiple coating layersfor any of the previously shown configurations.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (′TSFC)”—is the industry standardparameter of 1 bm of fuel being burned divided by 1 bf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram ° R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

Airfoils located downstream of combustor section 26, such as statorvanes and rotor blades in the turbine section 28, for example, operatein a high-temperature environment. The airfoils located in the highpressure turbine 54 are subjected to higher temperatures than airfoilslocated in the low pressure turbine 46. Airfoils that are exposed tohigh temperatures, such as those in the high pressure turbine 54,typically include cooling circuits with internal cooling channels andfilm cooling channels that direct a flow of cooling air through theairfoil to remove heat, reduce friction and mixing loss factors, andthus prolong the useful life of the airfoil. Airfoils in the lowpressure turbine are not actively cooled in this manner as they are notsubjected to temperatures as high as those experienced by airfoils inthe high pressure turbine 54.

By configuring these non-cooled low pressure turbine airfoils in themanner described below, efficiency can be improved and friction andmixing loss factors can be reduced. This efficiency increase is a directresult of configuring these airfoils such that a laminar boundary layercan be maintained as hot combustion gases flow along the externalsurface of the airfoils.

FIGS. 2-3 show an example of a turbine rotor blade 60 from the lowpressure turbine 46 which has a root section 62, a platform 64, and anairfoil body 66. Root section 62 is connected to a rotor in the lowpressure turbine 46 (FIG. 1) as known. The airfoil body 66 includes aleading edge 68, a trailing edge 70, a suction side wall 72, and apressure side wall 74. The airfoil body 66 extends from a base at theplatform 64 to a tip 76. The platform 64 connects the airfoil body 66 tothe root section 62. The leading edge 68, trailing edge 70, suction sidewall 72, and pressure side wall 74 extend outwardly away from theplatform 64 in a direction opposite from the root section 62. Suctionside wall 72 and pressure side wall 74 connect leading edge 68 andtrailing edge 70.

As shown in FIG. 2, the airfoil body 66 includes an internal cavity 78.This cavity 78 is not a cooling cavity or cooling passage. This cavity78 is typically formed in the airfoil bodies 66 of the low pressureturbine 46 as a weight savings measure. The airfoil body 66 ispositioned within a hot gas flow path 80 and has at least one inletopening 82 formed in an outer surface 84 of the airfoil body 66 todirect hot gas into the internal cavity 78. The airfoil body 66 alsoincludes at least one outlet opening 86 formed in the outer surface 84to expel the hot exhaust gas. The outlet opening 86 is located at alower pressure area in the internal cavity 78 than the at least oneinlet opening 82 such that hot gas is drawn into the internal cavity 78via the at least one inlet opening 82 and expelled out the at least oneoutlet opening 86 to maintain laminar boundary layer attachment alongthe outer surface 84 as hot gas flows against the airfoil body 66.

As shown in FIG. 2, the internal cavity 78 has an internal pressure Piand the hot gas flow path 80 has an external pressure P_(os). A pressuredifferential is maintained such that there is a sufficiently lowpressure level at the inlet to controllably ingest hot gas, incombination with a sufficiently high pressure level in the cavity 78 toexpel the hot gas via the outlet opening 86. As such, Pi is less than P∞at the inlet opening 82 and Pi is greater than P∞ at the outlet opening86. In one example, a pressure ratio at the inlet of P_(os)/Pi isapproximately 1.5-1.10, while the pressure ratio at the outlet ofP_(os)/Pi is approximately 0.90-0.95.

By drawing the hot exhaust gases into an upstream end of the internalcavity 78 and expelling the gases via a downstream end of the cavity 78,a laminar boundary layer can be maintained along the pressure and/orsuction sides of the airfoil body 66, which helps prevent stalled flowseparation. Essentially, this is comparable to film cooling but inreverse. Instead of drawing cooling air into the internal cavity to coolthe airfoil via film cooling hoes, hot exhaust gases are drawn into theinternal cavity via inlet openings to help maintain laminar flow.

In the example shown in FIGS. 2-3, the at least one inlet opening 82comprises a plurality of inlet openings 82 formed in one of or both ofthe pressure 74 and suction 72 side walls. The inlet openings 82 can beapertures, holes, slots, etc. and can have various shapes and/or sizes.Further, the leading edge 68 is free from inlet openings 82. In oneexample, a showerhead region 88 (identified by section line L in FIG. 2)of the airfoil body 66 does not include any inlet openings 82. Theshower head region 88 encompasses the leading edge 68 and transitionareas 88 a, 88 b that extend from the leading edge 68 along a shortdistance of the upstream portion of the pressure 74 and suction 72 sidewalls. The hot exhaust gas hits this showerhead region 88 which thendirects the gas along the sides of the body 66 where a portion of thegas is then ingested via the inlet openings 82, with a remaining portionof the gas continuing to flow along the sides of the airfoil body 66. Bynot having inlet openings 82 in the showerhead region 88, it prevents adirect input of a large amount of hot exhaust gas and instead allows amore controlled metering ingestion of the gas which allows the laminarboundary layer to be maintained.

FIG. 3 shows one example of the inlet openings 82 comprising slots. Inthis example the slots extend in a generally axial direction along thelength of the body 66. Each slot has a width W and a length L. In oneexample, the length L is longer than the width W. The lengths L andwidths W can be the same for each slot or variable. In one example, thespacing between the slots is on the order of 10:1 to 20:1, as the ratioof a distance D along the surface between slots to the slot width W.

In one example, the at least one outlet opening 86 is located near or atthe trailing edge 70 of the airfoil body 66. This location helps providethe desired transition from a lower pressure area of the internal cavity78 at the inlet to a lower pressure area at the outlet. The internalcavity 78 is wider at the leading edge end of the airfoil body andsubsequently narrows in a direction toward the trailing edge 70, whichresults in an increase in pressure needed to expel the hot exhaustgases. In another example, the at least one outlet opening 86 is locatednear or at the tip 76. This location utilizes centripetal pumping toassist in expelling the ingested hot exhaust flow.

As discussed above, the airfoil body 66 is not subjected to coolingairflow and does not include internal cooling passages. As such, theairfoil body 66 should be comprised of a material having a melting pointtemperature that is higher than a temperature of the hot gas flowingaround the airfoil body 66. Optionally, it is possible to achieve higheroperating temperature applicability for the component body by usingcoatings and thermal conductive enhancements.

FIGS. 4 and 5 are a comparison between flow conditions for a traditionalairfoil configuration (FIG. 4) and flow conditions for an airfoil bodyincorporating the subject invention (FIG. 5). Flow Un is directed towardthe leading edge 68 of the airfoil body. Each graph has a y-axis thatcorresponds to an increasing outer distance Δ from the external surface84 of the airfoil body and an x-axis that corresponds to an increasingdistance starting from the leading edge 68 and going towards thetrailing edge 70.

FIG. 4 is a flat plate nominal boundary layer diagram that shows aviscous sub-layer region 90 that is generally constant along the airfoilbody from the leading edge 68 to the trailing edge 70. A laminar regionof flow is indicated at 92 in a region that is just downstream from theleading edge 68. A transition region 94 shows the beginning of adeparture from laminar flow. This transition region 94 endsapproximately midway along the airfoil body 66 and then turns into aturbulent flow region 96 which becomes more turbulent in a direction ofthe trailing edge 70.

FIG. 5 is a flat plate enhanced laminar boundary layer diagram thatshows a similar viscous sub-layer region 90 that is generally constantalong the airfoil body from the leading edge 68 to the trailing edge 70;however, in this configuration there is no turbulent flow region. Alaminar region of flow is indicated at 98 in a region that is justdownstream from the leading edge 68, and as the flow starts to approachthe transition region of FIG. 4, the flow reaches at least one inletopening 82, which brings the flow back into a laminar boundary layer.This pattern continues down the length of the airfoil body such that theflow never moves out of the laminar flow region, i.e. turbulent flow isnot experienced by the airfoil body of FIG. 5.

Further, it should be noted that while the subject invention isdescribed as being used in an airfoil in a low pressure turbine, theinvention could also be used in other areas such as transition ducts orplatforms, for example. FIG. 6 shows an example of a transition duct 100as used in a mid-turbine frame 57 or turbine exhaust case. The duct 100includes inlet openings 102 that ingest gases in a manner that issimilar to that described above. FIG. 7 shows an example of platformlaminar flow where a platform 200 includes inlet openings 202 thatingest gases in a manner that is similar to that described above.

The subject invention utilizes hot gas inflow to maintain a laminarboundary layer along the pressure and/or suction sides of a componentbody 66, and to help prevent stalled flow separation. At least oneinternal cavity 78 of an otherwise uncooled blade, vane, or duct ismaintained at an appropriate pressure level, low enough to selectivelyand controllably ingest hot gas from the external gas path but stillhigh enough to expel the gas via an outlet from the component body 66.As discussed above, ingestion is through holes or thin radial slots withinternal metering features. This ingestion helps to maintain laminarboundary layer attachment and prevents its decomposition into aturbulent boundary layer and viscous sub-layer.

This concept has been successfully demonstrated in flight tests forusage on aircraft wings to decrease both aerodynamic drag and frictionalheating at high velocities. As a side effect, based on flat and conicalplate studies, a laminar boundary layer created on an airfoil reducesthe recovery factor towards a minimum of 0.81-0.83; otherwise therecovery factor approaching a more turbulent flow regime increasescloser to 0.90. Some of the benefits of the subject invention in thisconfiguration include the maintenance of laminar boundary layers in flowat extremely high Reynolds numbers (Re), to increase the turbine stageefficiency of an uncooled airfoil by as much as 1-2%. Further, with arecovery factor r=(Te−T)/(To/T) at typical LPT conditions (whereTe=Taw=adiabatic wall temperature; T=Tm=mean stream temperature; andTo=Ts=stagnation temperature), the lower recovery factor decreases theadiabatic wall temperature by a delta of as much as 14-15 degreesFahrenheit. The nominally lower adiabatic wall temperature can be tradedinto other factors such as durability, thrust and/or efficiency. Withthrust, increasing the LPT inlet temperature by 10-15 degrees Fahrenheitwill equal a +1% core thrust. For single crystal superalloys, a 14degree Fahrenheit increase in wall metal temperature yieldsapproximately 40% better creep life (by the trade factor 2{circumflexover ( )}(−ΔT/30° F.)). ΔT=delta metal temperature (degrees Fahrenheit).

In another example, this concept could be used to maintain boundarylayer attachment for compressor airfoils or fan blades in a mannersimilar to turbine airfoils. In this application there would be aeroefficiency and stall margin benefits. In one example, the compressorairfoil or fan blade is located within a flow path and would include atleast one inlet to an internal cavity as described above. With acompressor blade, the ingested air could be bled out of the internalcavity through an outlet via an attachment to a secondary air system.The universal application for this concept to work is the availabilityof a relative suction pressure inside the airfoil versus the outsideflow path.

In one example shown in FIG. 8, the airfoil body 66 is comprised of anon-metallic material. The transition duct 100 of FIG. 6 and theplatform 200 of FIG. 7 could also be made from a non-metallic material.One example of a non-metallic material is a ceramic matrix composite(CMC) material that has a high temperature capability. Examples of CMCmaterials include, but are not limited to: alumina matrix material,silicon, silica or silicon carbide materials and any variouscombinations thereof. The matrix can include embedded ceramic fiberssuch as oxidation stable reinforcing fibers including monofilaments likesilicon carbide for example, or can include yarn material comprisingalumina silicates, silicon carbide (NICALON®, SYLRAMIC®, etc. forexample), and/or chopped whiskers of similar materials. The CMC materialmay also include ceramic particles such oxides of Al, Si, Y, Zr and/orinorganic fillers as needed. It should be understood that these are justsome examples of CMC materials and that other types of CMC materialscould also be used. Optionally, monolithic ceramic or other similar highheat and highs strength materials could also be used.

In the example shown in FIG. 8, one or more of the inlet openings 82have passage surfaces that are coated with a coating 300 comprising atleast one of a thermal barrier coating (TBC) or environmental barriercoating (EBC). The inlet openings 102 in the duct 100 and the inletopenings 202 in the platform 200 may also include passage surfaces thatare coated with the coating 300 comprising at least one of a thermalbarrier coating or environmental barrier coating. As discussed above,high temperature exhaust flow is directed through the inlet openings 82and into the internal cavity 78. In one example, the surfaces of thecavity 78 and openings 82 can be exposed to temperatures as high as 2000degrees Celsius which can induce high thermal stress into the component.Further, the exhaust vapor exiting the combustor can degrade surfaces,such as a CMC material, for example. The TBC and/or EBC helps protectthese surfaces from experiencing high thermal stress and from degradingunder the high temperatures.

In one example, the EBC comprises a low conductivity coating that can becomprised of one or more coating layers. The EBC, for example, can havea thermal conductivity that is up to 10 times less than a thermalconductivity of the non-metallic material of the component body 66 toreduce the formation of thermal stresses in the body 66. In one example,the EBC coating comprises a first layer that includes silicon or silicaand at least one or more secondary layers that include any of thefollowing in any combination thereof: a rare earth silicate layer, ayttrium monosilicate, a mullite layer and/or a mullite and alkalineearth aluminosilicate layer, a barium strontium aluminosilicate layer,an ytterbium based layer, or layers of other similar materials.

As shown in the example of FIG. 8, the coating 300 is applied on atleast a portion of a laminar flow control aperture, e.g. an inletopening 82, of the body 66. As discussed above, the coating 300 may be aTBC 300 a, an EBC 300 b, or a combination of both as shown in FIG. 10.The coating 300 can be applied via chemical vapor deposition (CVD) orchemical vapor infiltration (CVI), or any other coating process.

In one example, all passage surfaces of the inlet openings 82 includethe coating 300. In one example, surfaces of all, or at least a portionof, the internal cavity 78 also include the coating 300, see FIG. 9 forexample. Additionally, as discussed above, the coatings 300 in this areacan be comprised of multiple layers of EBC and TBC in any order. In oneexample, the outermost layer of the coating 300 is comprised of a TBC.The coating 300 is especially useful when used in combination with acomponent body made from a CMC material. The coating 300 protects theexposed surfaces from any degrading effects of the combustor vapor gasesas the hot exhaust flow is directed through the inlet openings and drawninto the component internal cavity in a controlled manner such that thehot exhaust gas across the outer surface maintains a laminar boundarylayer. This coating 300 is also especially useful for LPT componentsthat do not receive a cooling airflow but instead control ingestion ofhot flow to maintain the laminar boundary layer to improve operatingefficiency.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

The invention claimed is:
 1. A gas turbine engine component comprising:a component body configured to be positioned within a flow path of a gasturbine engine having an external pressure, wherein the component bodycomprises an airfoil, and wherein the component body includes a leadingedge, a trailing edge, and pressure and suction side walls extendingfrom the leading edge to the trailing edge, and wherein the componentbody extends from a base to a tip to define a length and has a widthextending from the leading edge to the trailing edge; at least oneinternal cavity in the component body having an internal pressure, andwherein the component body is positioned downstream of a combustorsection and comprises a non-cooled component where the at least oneinternal cavity is free from receiving cooling flow; a plurality ofinlet openings formed in an outer surface of the component body todirect hot exhaust gas flow into the at least one internal cavity,wherein the leading edge is free from the plurality of inlet openings,and wherein the plurality of inlet openings comprise radial slots thatare spaced apart from each other along the width of the component bodyfrom the leading edge to the trailing edge and are spaced apart fromeach other along the length of the component body from the base to thetip; and at least one outlet from the at least one internal cavityformed at the trailing edge, wherein the internal pressure is less thanan inlet external pressure at the plurality of inlet openings and theinternal pressure is greater than an outlet external pressure at the atleast one outlet to controllably ingest hot exhaust gas via theplurality of inlet openings and expel the hot exhaust gas via the atleast one outlet to maintain a laminar boundary layer along the outersurface of the component body.
 2. The gas turbine engine componentaccording to claim 1 wherein the plurality of inlet openings are formedin both of the pressure and suction side walls.
 3. The gas turbineengine component according to claim 1 wherein the at least one outletcomprises at least one opening to the outer surface that is located nearor at the tip.
 4. The gas turbine engine component according to claim 1wherein the plurality of inlet openings provide a passage surface thatis coated with at least one of a thermal barrier coating orenvironmental barrier coating.
 5. The gas turbine engine componentaccording to claim 4 wherein the at least one internal cavity is coatedwith at least one of a thermal barrier coating or environmental barriercoating.
 6. The gas turbine engine component according to claim 5wherein the at least one internal cavity and the plurality of inletopenings are coated with a plurality of coatings including at least onethermal barrier coating and at least one environmental barrier coating.7. The gas turbine engine component according to claim 4 wherein thecomponent body is comprised of a non-metallic material.
 8. The gasturbine engine component according to claim 7 wherein the non-metallicmaterial is a ceramic matrix composite material.
 9. A gas turbine enginecomprising: a component body configured to be positioned within a flowpath of a gas turbine engine having an external pressure, wherein thecomponent body includes at least one internal cavity having an internalpressure, wherein the component body comprises an airfoil, and whereinthe component body includes a leading edge, a trailing edge, andpressure and suction side walls extending from the leading edge to thetrailing edge, and wherein the component body extends from a base to atip to define a length and has a width extending from the leading edgeto the trailing edge; a plurality of inlet openings formed in anexternal surface of the component body to direct hot exhaust gas flowinto the at least one internal cavity, wherein the plurality of inletopenings include a passage surface that is coated with at least one of athermal barrier coating or environmental barrier coating, and whereinthe plurality of inlet openings comprise radial slots that are spacedapart from each other along the width of the component body from theleading edge to the trailing edge and are spaced apart from each otheralong the length of the component body from the base to the tip; and atleast one outlet from the at least one internal cavity, wherein theinternal pressure is less than an inlet external pressure at theplurality of inlet openings and the internal pressure is greater than anoutlet external pressure at the at least one outlet to controllablyingest hot exhaust gas via the plurality of inlet openings and expel thehot exhaust gas via the at least one outlet to maintain a laminarboundary layer along the external surface of the component body.
 10. Thegas turbine engine according to claim 9 wherein the at least oneinternal cavity is free from receiving cooling air flow, and wherein theplurality of inlet openings are formed within the pressure and suctionside walls, and wherein the leading edge comprises a showerhead regionthat encompasses the leading edge and transition areas that extend fromthe leading edge along an upstream portion of the pressure and suctionside walls, and wherein the showerhead region is free from holes orslots, and wherein the plurality of inlet openings are spaced apart fromeach other along the pressure and suction side walls from downstream ofthe showerhead region to the trailing edge.
 11. The gas turbine engineaccording to claim 10 including: a compressor section; a combustorsection downstream of the compressor section; and a turbine sectiondownstream of the combustor section, and wherein the turbine sectionincludes at least a high pressure turbine downstream of the combustorsection and a low pressure turbine downstream of the high pressureturbine, and wherein the airfoil is located within the low pressureturbine.
 12. The gas turbine engine according to claim 9 wherein the atleast one internal cavity is coated with at least one of a thermalbarrier coating or environmental barrier coating.
 13. The gas turbineengine according to claim 9 wherein the at least one internal cavity andthe plurality of inlet openings are coated with a plurality of coatingsincluding at least one layer of a thermal barrier coating and at leastone layer of an environmental barrier coating.
 14. The gas turbineengine component according to claim 13 wherein the component body ismade of a non-metallic material comprising a ceramic matrix compositematerial.
 15. A method of enhancing laminar flow for a gas turbineengine component comprising the steps of: a) positioning a componentbody comprising an airfoil within a hot gas flow of a gas turbine enginehaving an external pressure, wherein the component body includes aleading edge, a trailing edge, and pressure and suction side wallsextending from the leading edge to the trailing edge, and wherein thecomponent body extends from a base to a tip to define a length and has awidth extending from the leading edge to the trailing edge, and whereinthe component body has at least one internal cavity having an internalpressure, and wherein the component body is positioned downstream of acombustor section; b) keeping the at least one internal cavity free fromcooling flow; c) providing a plurality of inlet openings formed in anexternal surface of the component body to direct hot exhaust gas flowinto the at least one internal cavity, wherein the plurality of inletopenings comprise radial slots that are spaced apart from each otheralong the width of the component body from the leading edge to thetrailing edge and are spaced apart from each other along the length ofthe component body from the base to the tip, and wherein the leadingedge is free from the plurality of inlet openings, and providing atleast one outlet from the at least one internal cavity to externalatmosphere, wherein the at least one outlet is at the trailing edge, andd) maintaining the internal pressure to be less than the externalpressure at the plurality of inlet openings and to be greater than theexternal pressure at the at least one outlet to controllably ingest aportion of the hot gas flow via the plurality of inlet openings andexpel ingested hot exhaust gas via the at least one outlet to maintain alaminar boundary layer of a remaining portion of the hot gas flow alongthe external surface of the component body.
 16. The method according toclaim 15 including forming the component body from a non-metallicmaterial, and coating at least one of the internal cavity and theplurality of inlet openings with at least one of a thermal barriercoating or environmental barrier coating.
 17. The method according toclaim 15 wherein the plurality of inlet openings are formed within thepressure and suction side walls, and wherein the airfoil includes ashowerhead region that encompasses the leading edge and transition areasthat extend from the leading edge along an upstream portion of thepressure and suction side walls, and wherein the showerhead region isfree from the plurality of inlet openings, and wherein the plurality ofinlet openings are spaced apart from each other along the pressure andsuction side walls from downstream of the showerhead region to thetrailing edge.
 18. The method according to claim 15 wherein spacingbetween adjacent radial slots is 10:1 to 20:1 as a ratio of a distancebetween adjacent radial slots along the external surface in a directionof an engine center longitudinal axis to a width of the radial slots.19. The gas turbine engine according to claim 9 wherein spacing betweenadjacent radial slots is 10:1 to 20:1 as a ratio of a distance betweenadjacent radial slots along the external surface in a direction of anengine center longitudinal axis to a width of the radial slots.
 20. Thegas turbine engine component according to claim 1 wherein the componentbody includes a showerhead region that encompasses the leading edge andtransition areas that extend from the leading edge along an upstreamportion of pressure and suction side walls, and wherein the showerheadregion is free from the plurality of inlet openings, and wherein theradial slots each have a slot height that is greater than a slot width,and wherein the radial slots are spaced apart from each other along thewidth of the pressure and suction side walls from downstream of theshowerhead region to the trailing edge, and are spaced apart from eachother along the length from the base to the tip.
 21. The gas turbineengine component according to claim 1 wherein spacing between adjacentradial slots is 10:1 to 20:1 as a ratio of a distance between adjacentradial slots along the outer surface in a direction of an engine centerlongitudinal axis to a width of the radial slots.